Open cycle liquid rocket engine thrust chamber torch low pressure ignition device and method

Abstract

The present invention relates to a low-pressure ignition device and method for a flare of an open-cycle liquid rocket engine thrust chamber. The ignition device comprises a cylindrical ignition chamber, an ignition fuel delivery system and an ignition oxidant delivery system. The ignition chamber is used to be arranged in the thrust chamber of the rocket engine. One end of the ignition chamber is open, and the other end is provided with an igniter. The outer surface of the ignition chamber is provided with a fuel flow channel and an oxidant flow channel along the axial direction. The inner wall of the other end of the ignition chamber is provided with at least one group of fuel spray holes and oxidant spray holes. The ignition fuel delivery system is connected to one end of the fuel flow channel, and the fuel spray hole is connected to the other end of the fuel flow channel. The ignition oxidant delivery system is connected to one end of the oxidant flow channel, and the oxidant spray hole is connected to the other end of the oxidant flow channel. Advantages: improve the ignition reliability of the liquid rocket engine, simplify the complexity of the ignition system, improve the reliability of the ignition process of the liquid rocket engine, and simplify the system composition, improve the reliability of the system, and reduce the cost of the system.

Description

Technical Field

[0001] The invention relates to the technical field of rocket engines, and in particular to a low-pressure ignition device and method for a flare in a thrust chamber of an open-cycle liquid rocket engine. Background Art

[0002] The common power cycles of liquid rocket engines include gas generator cycle, expansion cycle and regenerative cycle. The gas generator cycle is a typical open cycle, while the expansion cycle and regenerative cycle are typical closed cycles. Among them, the gas generator cycle is generally called an open cycle because the gas used to do work in the turbine is not returned to the mainstream for combustion. The expansion cycle and regenerative cycle are called closed cycles because the medium (propellant or gas) used to do work in the turbine eventually returns to the mainstream and is burned and ejected together. Open cycle liquid rocket engines are widely used in various space launch missions, and the ignition process of their thrust chambers is a key step to ensure the normal operation of the engine.

[0003] At present, common ignition methods include gunpowder ignition, electric spark ignition, hot gas ignition, etc. However, the existing technology still has certain deficiencies in ignition reliability and the life of the ignition device, which affects the overall performance of the rocket engine. Therefore, it is urgent to develop an efficient and reliable ignition method. The existing technology has carried out research and experiments on the low-pressure torch-type electric ignition system of the expansion cycle hydrogen-oxygen engine (Li Jinjiang et al. "Research on the Torch Ignition System Scheme of the Expansion Cycle Hydrogen-Oxygen Engine" Missile and Space Launch Technology 2024), which draws hydrogen and oxygen media from the front of the valve, relies on spark plug ignition to burn small streams of hydrogen and oxygen to form a torch, and then ignites the entire thrust chamber. However, the existing technology is based on the expansion cycle hydrogen-oxygen rocket engine, which is limited by the power cycle mode of the expansion cycle (closed cycle) and the propellant medium of hydrogen and oxygen. It has poor adaptability and high difficulty in promotion on the liquid oxygen-methane engine of the gas generator cycle (open cycle).

[0004] Based on this, it is necessary to develop a low-pressure ignition device and method for the thrust chamber torch of an open-cycle liquid rocket engine to overcome the above-mentioned technical problems. Summary of the invention

[0005] The technical problem to be solved by the present invention is to provide a low-pressure ignition device and method for a thrust chamber torch of an open-cycle liquid rocket engine, which effectively overcomes the defects of the prior art.

[0006] The technical solution of the present invention to solve the above technical problems is as follows:

[0007] A low-pressure ignition device for a thrust chamber torch of an open-cycle liquid rocket engine comprises a cylindrical ignition chamber, an ignition fuel delivery system and an ignition oxidant delivery system, wherein the ignition chamber is used to be arranged in the thrust chamber of a rocket engine, one end of the ignition chamber is open, and the other end is provided with an igniter, a fuel flow channel and an oxidant flow channel are axially provided on the outer surface of the ignition chamber, and at least one group of fuel nozzles and oxidant nozzles are provided on the inner wall of the other end of the ignition chamber, the ignition fuel delivery system is connected to one end of the fuel flow channel, the fuel nozzle is connected to the other end of the fuel flow channel, the ignition oxidant delivery system is connected to one end of the oxidant flow channel, and the oxidant nozzle is connected to the other end of the oxidant flow channel.

[0008] Based on the above technical solution, the present invention can also be improved as follows.

[0009] Furthermore, the above-mentioned fuel flow channels and oxidant flow channels are respectively provided with a plurality of them, and are staggered and distributed along the circumference of the above-mentioned ignition chamber, and the above-mentioned fuel flow channels and oxidant flow channels respectively extend along the axial direction of the above-mentioned ignition chamber, one end of all the above-mentioned fuel flow channels is commonly connected to a circular first diversion chamber, one end of all the above-mentioned oxidant flow channels is commonly connected to a circular second diversion chamber, the above-mentioned ignition fuel delivery system is connected to the above-mentioned first diversion chamber, and the above-mentioned ignition oxidant delivery system is connected to the above-mentioned second diversion chamber.

[0010] Furthermore, the injection directions of the fuel injection holes and the oxidant injection holes are all along the tangent direction of the inner wall of the ignition chamber, and the injection directions of the same group of the fuel injection holes and the oxidant injection holes are relatively close.

[0011] Furthermore, the fuel injection holes and the oxidant injection holes are provided in two or three groups, and a central angle formed by the fuel injection holes and the oxidant injection holes in the same group in the circumferential direction is less than 180°.

[0012] Furthermore, the igniter is a pneumatic resonance igniter or an electric spark igniter.

[0013] Furthermore, the ignition fuel delivery system includes an ignition fuel inlet conduit, an ignition fuel outlet conduit and a first electrically controlled valve, one end of the ignition fuel inlet conduit is connected to a fuel source, and the other end is connected to an inlet of the first electrically controlled valve, one end of the ignition fuel outlet conduit is connected to an outlet of the first electrically controlled valve, and the other end is connected to the fuel flow channel.

[0014] Furthermore, the above fuel is methane.

[0015] Furthermore, the above-mentioned ignition oxidant delivery system includes an ignition oxidant inlet conduit, an ignition oxidant outlet conduit and a second electrically controlled valve, one end of the above-mentioned ignition oxidant inlet conduit is connected to the liquid oxidant, and the other end is connected to the inlet of the above-mentioned second electrically controlled valve, one end of the above-mentioned ignition oxidant outlet conduit is connected to the outlet of the above-mentioned second electrically controlled valve, and the other end is connected to the above-mentioned oxidant flow channel.

[0016] Furthermore, the liquid oxidant is liquid oxygen.

[0017] The beneficial effects of the present invention are: improving the ignition reliability of the liquid rocket engine, simplifying the complexity of the ignition system, improving the reliability of the ignition process of the liquid rocket engine, and at the same time simplifying the system composition, improving the system reliability and reducing the system cost.

[0018] A method for igniting a flare in a thrust chamber of an open-cycle liquid rocket engine is also provided, which is implemented by using a low-pressure ignition device for igniting a flare in a thrust chamber of an open-cycle liquid rocket engine, and comprises the following steps:

[0019] Step 1: Turn on the igniter. At the same time, the rocket engine starts to deliver fuel and oxidizer to the thrust chamber, and delivers fuel and oxidizer to the ignition chamber through the ignition fuel delivery system and the ignition oxidizer delivery system;

[0020] Step 2: Ignite the mixed fuel and oxidant in the ignition chamber through the igniter, and form a torch at one end of the ignition chamber, and spray it into the main combustion chamber of the thrust chamber. The high-temperature and high-pressure ignition torch flame triggers the combustion of the fuel and oxidant mixture in the combustion chamber. After the thrust chamber is successfully ignited, the ignition oxidant delivery system is cut off, and the ignition fuel delivery system continues to supply ignition fuel to protect the structure;

[0021] Step 3: The fully burned fuel and oxidizer in the main combustion chamber of the thrust chamber are ejected through the nozzle structure of the thrust chamber, and the overall thrust of the engine is established.

[0022] The beneficial effects are: the method can significantly improve the reliability of the ignition device, ensure the successful and effective ignition of the thrust chamber, reduce the ignition delay, and increase the number of times the ignition device is reused. BRIEF DESCRIPTION OF THE DRAWINGS

[0023] Figure 1 The present invention is a schematic structural diagram of the low-pressure ignition device for the thrust chamber torch of an open-cycle liquid rocket engine.

[0024] In the accompanying drawings, the components represented by the reference numerals are listed as follows:

[0025] 1. Ignition chamber; 2. Ignition fuel delivery system; 3. Ignition oxidant delivery system; 4. Ignitor; 11. Fuel flow channel; 12. Oxidant flow channel; 13. Fuel spray hole; 14. Oxidant spray hole; 15. First diverter cavity; 16. Second diverter cavity; 21. Ignition fuel inlet conduit; 22. Ignition fuel outlet conduit; 23. First electrically controlled valve; 31. Ignition oxidant inlet conduit; 32. Ignition oxidant outlet conduit; 33. Second electrically controlled valve. DETAILED DESCRIPTION

[0026] The principles and features of the present invention are described below in conjunction with the accompanying drawings. The examples given are only used to explain the present invention and are not used to limit the scope of the present invention.

[0027] Example: Figure 1 As shown, the open-cycle liquid rocket engine thrust chamber torch low-pressure ignition device of this embodiment includes a cylindrical ignition chamber 1, an ignition fuel delivery system 2 and an ignition oxidant delivery system 3, the ignition chamber 1 is used to be arranged in the thrust chamber of the rocket engine, one end of the ignition chamber 1 is open, and the other end is provided with an igniter 4, the outer surface of the ignition chamber 1 is axially provided with a fuel flow channel 11 and an oxidant flow channel 12, the inner wall of the other end of the ignition chamber 1 is provided with at least one group of fuel nozzles 13 and oxidant nozzles 14, the ignition fuel delivery system 2 is connected to one end of the fuel flow channel 11, the fuel nozzle 13 is connected to the other end of the fuel flow channel 11, the ignition oxidant delivery system 3 is connected to one end of the oxidant flow channel 12, and the oxidant nozzle 14 is connected to the other end of the oxidant flow channel 12.

[0028] The method for implementing ignition using the open cycle liquid rocket engine thrust chamber torch low-pressure ignition device of this embodiment comprises the following steps:

[0029] Step 1: Turn on the igniter 4. At the same time, the rocket engine starts to deliver fuel and oxidant to the thrust chamber, and delivers fuel and oxidant to the ignition chamber 1 through the ignition fuel delivery system 2 and the ignition oxidant delivery system 3;

[0030] Step 2: The mixed fuel and oxidant in the ignition chamber 1 are ignited by the igniter 4, and a torch is formed at one end of the ignition chamber 1, and sprayed into the main combustion chamber of the thrust chamber. The high-temperature and high-pressure ignition torch flame triggers the combustion of the fuel and oxidant mixture in the combustion chamber. After the thrust chamber is successfully ignited, the ignition oxidant delivery system 3 is cut off, and the ignition fuel delivery system 2 keeps supplying the ignition fuel to protect the structure.

[0031] Step 3: The fully burned fuel and oxidizer in the main combustion chamber of the thrust chamber are ejected through the nozzle structure of the thrust chamber, and the overall thrust of the engine is established.

[0032] In this embodiment, the above-mentioned fuel flow channels 11 and the oxidant flow channels 12 are respectively provided with a plurality of them, and are staggered and distributed along the circumference of the above-mentioned ignition chamber 1, and the above-mentioned fuel flow channels 11 and the oxidant flow channels 12 respectively extend along the axial direction of the above-mentioned ignition chamber 1, one end of all the above-mentioned fuel flow channels 11 is commonly connected to a circular first diversion chamber 15, one end of all the above-mentioned oxidant flow channels 12 is commonly connected to a circular second diversion chamber 16, the above-mentioned ignition fuel delivery system 2 is connected to the above-mentioned first diversion chamber 15, and the above-mentioned ignition oxidant delivery system 3 is connected to the above-mentioned second diversion chamber 16.

[0033] More specifically, the ignition chamber 1, the fuel flow channel 11 and the oxidant flow channel 12 are manufactured based on 3D printing technology, wherein there are a total of 12 fuel flow channels 11 and oxidant flow channels 12, which are staggered and spaced circumferentially and numbered 1#, 2#, 3#...12#, wherein odd-numbered ones such as 1#, 3#, 5#...11# are fuel flow channels 11, which help to heat up the ignition methane, and even-numbered ones such as 2#, 4#, 6#...12# are oxidant flow channels 12, which help to heat up the ignition oxygen, so as to achieve heating of the ignition methane and the ignition oxygen, so as to form a gaseous mixture in the ignition chamber 1.

[0034] As a preferred embodiment, the injection directions of the fuel injection holes 13 and the oxidant injection holes 14 are all along the tangent direction of the inner wall of the ignition chamber 1, and the injection directions of the same group of the fuel injection holes 13 and the oxidant injection holes 14 are relatively close.

[0035] In the above embodiment, the fuel injection hole 13 and the oxidant injection hole 14 are both arranged along the tangent direction of the ignition chamber 1 and enter the ignition chamber 1 in the direction of the swirl flow, which can promote the full mixing of the gaseous fuel and the gaseous oxidant.

[0036] As a preferred embodiment, the fuel injection holes 13 and the oxidant injection holes 14 are provided in two or three groups, and the central angle formed by the fuel injection holes 13 and the oxidant injection holes 14 in the same group in the circumferential direction is less than 180°.

[0037] In the above embodiment, the fuel spray holes 13 and the oxidant spray holes 14 are provided in two or three groups, each group has one fuel spray hole 13 and one oxidant spray hole 14, and the central angle formed by the fuel spray holes 13 and the oxidant spray holes 14 in each group in the circumferential direction is less than 180°, and the spray directions are opposite, which makes the fuel and the oxidant to be oppositely injected (that is, the fuel enters in a clockwise swirl, and the oxidant enters in a counterclockwise swirl), and better mixed.

[0038] In this embodiment, the igniter 4 is a pneumatic resonance igniter or an electric spark igniter. It should be noted that before the ignition fuel and the ignition oxidant flow in normally, the igniter 4 has formed a stable ignition state to ensure a smooth initial ignition process.

[0039] As a preferred embodiment, the ignition fuel delivery system 2 includes an ignition fuel inlet conduit 21, an ignition fuel outlet conduit 22 and a first electrically controlled valve 23. One end of the ignition fuel inlet conduit 21 is connected to a fuel source (that is, connected to the thrust chamber regeneration cooling outlet), and the other end is connected to the inlet of the first electrically controlled valve 23. One end of the ignition fuel outlet conduit 22 is connected to the outlet of the first electrically controlled valve 23, and the other end is connected to the fuel flow channel 11.

[0040] In the above-mentioned implementation scheme, the fuel flows to the fuel flow channel 11 through the ignition fuel inlet conduit 21 and the ignition fuel outlet conduit 22. During this period, the flow rate is controlled by the first electrically controlled valve 23. In addition, during the flow process, heat is mutually conducted between the fuel and the ignition fuel inlet conduit 21, the ignition fuel outlet conduit 22 and the fuel flow channel 11, and the fuel is heated, so that the ignition fuel injected into the ignition chamber 1 is a normalized gaseous medium.

[0041] In this embodiment, the fuel is methane.

[0042] As a preferred embodiment, the ignition oxidant delivery system 3 includes an ignition oxidant inlet conduit 31, an ignition oxidant outlet conduit 32 and a second electrically controlled valve 33. One end of the ignition oxidant inlet conduit 31 is connected to the liquid oxidant (that is, connected to the outlet of the thrust chamber oxygen valve), and the other end is connected to the inlet of the second electrically controlled valve 33. One end of the ignition oxidant outlet conduit 32 is connected to the outlet of the second electrically controlled valve 33, and the other end is connected to the oxidant flow channel 12.

[0043] In the above-mentioned embodiment, the liquid oxidant enters the oxidant flow channel 12 through the ignition oxidant inlet conduit 31 and the ignition oxidant outlet conduit 32, and heat is mutually conducted between the ignition oxidant inlet conduit 31, the ignition oxidant outlet conduit 32 and the oxidant flow channel 12, so that the liquid oxidant is heated up, so that the ignition oxidant injected into the ignition chamber 1 is a gaseous medium.

[0044] In this embodiment, the liquid oxidant is liquid oxygen.

[0045] It should be noted that before ignition, the ignition fuel delivery system 2 and the ignition oxidant delivery system 3 deliver fuel and oxidant to the ignition chamber 1. When the fuel and oxidant pass through their respective corresponding pipelines and fuel flow channels 11 and oxidant flow channels 12, the structural heat sinks of the pipelines and flow channels themselves are used to vaporize the fuel and oxidant that enter first to form normal temperature (not room temperature) gas and oxygen, which are filled into the ignition chamber 1 of the ignition device. After subsequent ignition to form stable combustion, the heat generated by the combustion is used to heat the fuel and oxidant, and finally they are vaporized and enter the ignition chamber 1. When the combustion gas ejected from the ignition chamber 1 successfully ignites the thrust chamber, stable combustion is formed in the thrust chamber, and then the oxidant delivery is cut off. More specifically, in order to ensure the reliable operation of the ignition device and coordinate with the operation of the thrust chamber, during the ignition period, the ignition time is at least 2s. After the ignition is completed, the first electrically controlled valve 23 is controlled to cut off the oxidant supply to the ignition chamber 1, and the ignition chamber 1 is switched to fuel supply only. A small amount of liquid methane is continuously supplied to the thrust chamber through the ignition chamber 1, forming an air seal for the ignition chamber 1 while reliably cooling the portion where the ignition device is connected to the high-temperature combustion gas of the thrust chamber to ensure that the structure therein is not corroded by the high-temperature combustion gas.

[0046] The flare ignition system contained in the present invention does not require additional blowing settings. The blowing and cleaning of the system is carried out together with the blowing and cleaning of the thrust chamber. When the engine performs regenerative cooling of the thrust chamber and blowing of the oxygen valve outlet, the flare ignition system opens the electronically controlled valves (the first electronically controlled valve and the second electronically controlled valve) to ensure that the flow paths of the ignition fuel path and the ignition oxidant path are unobstructed, and the blowing medium of the original engine thrust chamber is used to achieve the purging of the flare ignition system. In addition, the flare ignition system of this embodiment does not require additional temperature recovery settings. After the flare ignition system is finished working, since the continuously supplied methane may still be in a low temperature or extremely low temperature state, the flare ignition system structure itself is still in a low temperature state. After the whole engine is finished working, the fuel of the ignition system is naturally cut off, and the ignition system and the surrounding structures conduct heat conduction to each other and return to normal temperature together.

[0047] In the description of the present invention, it is to be understood that the terms “center”, “longitudinal”, “lateral”, “length”, “width”, “thickness”, “up”, “down”, “front”, “back”, “left”, “right”, “vertical”, “horizontal”, “top”, “bottom”, “inside”, “outside”, “clockwise”, “counterclockwise”, “axial”, “radial”, “circumferential”, etc., indicating orientations or positional relationships based on the orientations or positional relationships shown in the accompanying drawings, are only for the convenience of describing the present invention and simplifying the description, and do not indicate or imply that the referred device or element must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be understood as limiting the present invention.

[0048] In addition, the terms "first" and "second" are used for descriptive purposes only and should not be understood as indicating or implying relative importance or implicitly indicating the number of the indicated technical features. Therefore, the features defined as "first" and "second" may explicitly or implicitly include at least one of the features. In the description of the present invention, the meaning of "plurality" is at least two, such as two, three, etc., unless otherwise clearly and specifically defined.

[0049] In the present invention, unless otherwise clearly specified and limited, the terms "installed", "connected", "connected", "fixed" and the like should be understood in a broad sense, for example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection or an indirect connection through an intermediate medium, it can be the internal connection of two elements or the interaction relationship between two elements, unless otherwise clearly defined. For ordinary technicians in this field, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0050] In the present invention, unless otherwise clearly specified and limited, a first feature being "above" or "below" a second feature may mean that the first and second features are in direct contact, or the first and second features are in indirect contact through an intermediate medium. Moreover, a first feature being "above", "above" or "above" a second feature may mean that the first feature is directly above or obliquely above the second feature, or simply means that the first feature is higher in level than the second feature. A first feature being "below", "below" or "below" a second feature may mean that the first feature is directly below or obliquely below the second feature, or simply means that the first feature is lower in level than the second feature.

[0051] In the description of this specification, the description with reference to the terms "one embodiment", "some embodiments", "example", "specific example", or "some examples" etc. means that the specific features, structures, materials or characteristics described in conjunction with the embodiment or example are included in at least one embodiment or example of the present invention. In this specification, the schematic representations of the above terms do not necessarily refer to the same embodiment or example. Moreover, the specific features, structures, materials or characteristics described may be combined in any one or more embodiments or examples in a suitable manner. In addition, those skilled in the art may combine and combine the different embodiments or examples described in this specification and the features of the different embodiments or examples, without contradiction.

[0052] Although the embodiments of the present invention have been shown and described above, it is to be understood that the above embodiments are exemplary and are not to be construed as limitations of the present invention. A person skilled in the art may change, modify, replace and vary the above embodiments within the scope of the present invention.

Claims

1. A low-pressure ignition device for a thrust chamber torch of an open-cycle liquid rocket engine, characterized in that: The invention comprises a cylindrical ignition chamber (1), an ignition fuel delivery system (2) and an ignition oxidant delivery system (3), wherein the ignition chamber (1) is used to be arranged in the thrust chamber of a rocket engine, the ignition chamber (1) is open at one end and is provided with an igniter (4) at the other end, the outer surface of the ignition chamber (1) is provided with a fuel flow channel (11) and an oxidant flow channel (12) along the axial direction, the inner wall of the other end of the ignition chamber (1) is provided with at least one group of fuel injection holes (13) and oxidant injection holes (14), the ignition fuel delivery system (2) is connected to one end of the fuel flow channel (11), the fuel injection holes (13) are connected to the other end of the fuel flow channel (11), the ignition oxidant delivery system (3) is connected to one end of the oxidant flow channel (12), the oxidant injection holes (14) are connected to the other end of the oxidant flow channel (12), and the fuel flow channel (11) is connected to the inner wall of the other end of the ignition chamber (1). A plurality of fuel flow channels (11) and oxidant flow channels (12) are provided respectively, and are staggered and spaced along the circumference of the ignition chamber (1), and the fuel flow channels (11) and oxidant flow channels (12) respectively extend along the axial direction of the ignition chamber (1), one end of all the fuel flow channels (11) are connected to a first annular flow diversion chamber (15), one end of all the oxidant flow channels (12) are connected to a second annular flow diversion chamber (16), the ignition fuel delivery system (2) is connected to the first flow diversion chamber (15), the ignition oxidant delivery system (3) is connected to the second flow diversion chamber (16), the injection directions of the fuel injection holes (13) and the oxidant injection holes (14) are both along the tangent direction of the inner wall of the ignition chamber (1), and the injection directions of the fuel injection holes (13) and the oxidant injection holes (14) of the same group are relatively close; The ignition method comprises the following steps: Step 1: Turn on the igniter (4), and at the same time, the rocket engine starts to deliver fuel and oxidant to the thrust chamber, and delivers fuel and oxidant to the ignition chamber (1) through the ignition fuel delivery system (2) and the ignition oxidant delivery system (3); Step 2: The mixed fuel and oxidant in the ignition chamber are ignited by the igniter (4), and a torch is formed at one end of the ignition chamber (1) and sprayed into the main combustion chamber of the thrust chamber. The high-temperature and high-pressure ignition torch flame triggers the combustion of the fuel and oxidant mixture in the combustion chamber. After the thrust chamber is successfully ignited, the ignition oxidant delivery system (3) is cut off, and the ignition fuel delivery system (2) continues to supply the ignition fuel to protect the structure. Step 3: The fully burned fuel and oxidizer in the main combustion chamber of the thrust chamber are ejected through the nozzle structure of the thrust chamber, and the overall thrust of the engine is established.

2. The low-pressure ignition device for the thrust chamber torch of an open-cycle liquid rocket engine according to claim 1, characterized in that: The fuel injection holes (13) and the oxidant injection holes (14) are provided in two or three groups, and the central angle formed by the fuel injection holes (13) and the oxidant injection holes (14) in the same group in the circumferential direction is less than 180°.

3. The low-pressure ignition device for the thrust chamber torch of an open-cycle liquid rocket engine according to claim 1, characterized in that: The igniter (4) is a pneumatic resonance igniter or an electric spark igniter.

4. The low-pressure ignition device for the thrust chamber torch of an open-cycle liquid rocket engine according to claim 1, characterized in that: The ignition fuel delivery system (2) comprises an ignition fuel inlet conduit (21), an ignition fuel outlet conduit (22) and a first electrically controlled valve (23); one end of the ignition fuel inlet conduit (21) is connected to a fuel source, and the other end is connected to an inlet of the first electrically controlled valve (23); one end of the ignition fuel outlet conduit (22) is connected to an outlet of the first electrically controlled valve (23), and the other end is connected to the fuel flow channel (11).

5. The low-pressure ignition device for the thrust chamber torch of an open-cycle liquid rocket engine according to claim 4, characterized in that: The fuel is methane.

6. The low-pressure ignition device for the thrust chamber torch of an open-cycle liquid rocket engine according to claim 1, characterized in that: The ignition oxidant delivery system (3) comprises an ignition oxidant inlet conduit (31), an ignition oxidant outlet conduit (32) and a second electrically controlled valve (33); one end of the ignition oxidant inlet conduit (31) is connected to the liquid oxidant, and the other end is connected to the inlet of the second electrically controlled valve (33); one end of the ignition oxidant outlet conduit (32) is connected to the outlet of the second electrically controlled valve (33), and the other end is connected to the oxidant flow channel (12).

7. The low-pressure ignition device for the thrust chamber torch of an open-cycle liquid rocket engine according to claim 6, characterized in that: The liquid oxidant is liquid oxygen.

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